The present invention relates generally to gas turbine engines. More particularly, the present invention relates to a combustor liner support assembly for disposition within a gas turbine engine. Although the present invention was developed for use in a gas turbine engine, certain applications of the invention may fall outside of this field.
A gas turbine engine is typical of the type of turbo machinery in which the present invention described herein may be advantageously employed. It is well known that a gas turbine engine conventionally comprises a compressor for compressing inlet air to an increased pressure. A mixture of fuel and increased pressure air is ignited and burned in the combustor to generate a high temperature gaseous flow stream which is discharged from the combustor and directed by a series of turbine vanes to a series of turbine blades. The turbine blades convert the thermal energy from the high temperature gaseous flow stream into rotational kinetic energy, which in turn drives the compressor and a propeller, fan or other similar device. Alternatively, the high temperature gaseous flow stream may be used directly as a thrust for providing motive force, such as in a turbine jet engine.
In a typical gas turbine engine, relatively low pressure combustor gases flow through the interior of an annular combustor liner. Higher pressure compressor discharge cooling gas flows across the exterior of the combustor liner to lower the temperature of the liner by way of conventional film cooling and/or impingement cooling. The combustor liner is commonly supported at its upstream end, and its downstream end is allowed to float relative to the first turbine vane or nozzle. In such a support arrangement, a seal is sometimes provided in the area between the downstream end of the combustor liner and the turbine nozzle to restrict passage of the high pressure cooling gas into the flow path of the low-pressure combustor gases. However, different thermal coefficients of expansion between the combustor liner and the turbine nozzle results in increased sealing difficulties in this area of the gas turbine engine. More specifically, various components of the combustor and turbine sections thermally expand and contract at different rates. Thus, a seal disposed between the combustor and turbine sections must accommodate for thermal expansion and contraction to avoid, or at least minimize, damage to the combustor liner or the first turbine vane and prevent an unacceptably short useful life thereof.
Additionally, during the operation of the gas turbine engine, the combustor liner serves as the inner wall of a pressure vessel that must withstand the pressure differential between the high pressure cooling gas and the low pressure combustor gases. This pressure differential imposes a buckling load on the combustor liner, which in turn must be suitably configured to resist inward deflection. This is sometimes accomplished, for example, by increasing the thickness of the liner to correspondingly increase its hoop strength. However, this usually creates undesirable higher thermal gradients and hoop stresses within the liner, which can lead to low cycle fatigue (LCF), possibly resulting in liner failure or a shortened useful liner life. Alternatively, an annular support having a high moment of inertia is sometimes attached directly to the outside of the combustor liner to provide sufficient support to resist buckling. However, different thermal coefficients of expansion between the liner and the annular support again leads to undesirable higher thermal gradients and stresses within the liner.
Heretofore, there has been a need for a combustor liner support assembly which accommodates for the thermal cycle operation inherent in a gas turbine engine. An effective means for satisfying this need has escaped those skilled in the art. The present invention satisfies this need in a novel and unobvious way.
One form of the present invention contemplates a combustor liner support assembly for a gas turbine engine.
In another form of the present invention, a gas turbine engine includes an annular combustor having a combustor liner, and a plurality of circumferentially spaced apart projections extending from the combustor liner. A ring is disposed about the combustor liner and positioned adjacent the projections. Each of the projections engages a corresponding circumferentially facing portion of the ring and circumferentially supports the combustor liner while allowing relative radial displacement between the combustor liner and the ring.
In a further form, an apparatus is provided for supporting an annular combustor liner in a gas turbine engine. The apparatus includes a number of lugs spaced circumferentially about the combustor liner and extending from an outer surface of the combustor liner. A band is disposed about the combustor liner and is coupled to the lugs. One of the band and the lugs defines a plurality of radial slots, while the other defines a complementary portion disposed within each of the radial slots to provide circumferential support to the combustor liner while allowing relative radial movement between the combustor liner and the band.
In still another form, a combustor support assembly in a gas turbine engine includes a combustor liner extending along a flow path. A support member is disposed about the periphery of the combustor liner and is coupled thereto. One of the combustor liner and the support member defines a plurality of projections spaced about the periphery of the combustor liner, while the other defines a plurality of pairs of opposing surfaces extending in a transverse direction and correspondingly spaced about the periphery of the combustor liner. Each of the projections is disposed between a corresponding pair of opposing surfaces and bears against at least one of the opposing bearing surfaces to provide peripheral support to the combustor liner while being displaceable between the opposing surfaces to allow relative movement between the combustor liner and the support member in a transverse direction.
Further forms, embodiments, objects, features, advantages, benefits, and aspects of the present invention shall become apparent from the drawings and descriptions provided herein.